A significant advantage of turboprop engines in comparison to, for example, turbofan engines is that the turboprop engine can operate in the cruise mode at a relatively low specific fuel consumption. However, there are with the turboprop engine certain features that result in design problems which are of less significance with respect to turbofan engines In a modern turbofan engine having a bypass ratio as high as possibly five or six, the diameter of the fan is of a size that it can be driven directly from the turbine at the same rotational speed as the turbine not requiring a gear box or transmission. Also, the rotating blades of the fan generally have a fixed pitch, while in a turboprop engine there are generally variable pitch blades. Further, in a turbofan engine, it is possible to match the compressor and turbine to optimize specific fuel consumption for the cruise mode, with this same match working quite well in the take-off and climb mode of operation.
On the other hand, in a turboprop engine, the bypass ratio of the propeller may be as high as eighty. Because of the large propeller diameter and relatively small engine size, the rotational speed of the propeller must be reduced, relative to the rotational speed of the turbine, this being accomplished by placing a speed reducing gear drive between the turbine shaft and the propeller. Further, to optimize operation of the turboprop engine, the blades must be mounted for variable pitch for two reasons. First, during the various operating modes of the engine, the pitch (i.e. angular position) of the blades should be changed to obtain optimum efficiency. Second, it is desirable that the propeller blades be able to be moved to a position to provide reverse thrust, thus eliminating the need for the thrust reverser of a conventional turbofan engine. For a turboprop engine that is to be used in a present day twin engine, 120 passenger commercial aircraft flying near Mach 0.8, the weight of the speed reducing gear transmission in combination with the apparatus to provide variable pitch to the propeller blades may be as high as 1000 pounds or more (i.e. 454 kg or more).
With regard to the matching of engine components, this becomes a problem of substantially greater significance in comparison with turbofan engines, primarily because of the substantially greater bypass ratio typical of turboprop engines. When the turboprop engine is operating in the cruise mode at a higher altitude, the engine operates most efficiently with the propeller blades set at a relatively steep pitch, and with the compressor operating at a relatively high compression ratio (i.e. in the order of 30 to one or possibly as high as 40 to one) or possibly higher for future designs. However, with the engine components sized and matched to operate at relatively low specific fuel consumption in the cruise mode, this same engine is overpowered for the take-off and climb mode of operation. During take-off and climb, the propeller blades operate most effectively when they are set at a relatively shallow pitch and are rotated at about the same speed as during the cruise mode. In these circumstances, the propeller blade does not need the full power which the turbine can generate during take-off and climb conditions when operating in a range of reasonable efficiency. If the engine were operated at such a power setting during take-off and climb, this excess power would be passed to the propeller blades, which in turn would have to be set at a steeper pitch to absorb the power generated by the turbine. Not only would this cause the propeller blades to operate more inefficiently, but it would substantially increase the torque loads transmitted through the gear box. This would in turn either overstress the components of the gear transmission or require the gear box to be designed for such higher loads, this resulting in unneccessary increased weight.
Accordingly, the more common prior art approach has been to either throttle the engine back during take-off and climb or derate the engine in some manner for this mode. However, throttling the engine back simply causes the engine to operate in a condition which is in the more inefficient range of the specific fuel consumption curve. Derating the engine, for example by bleeding air for the compressor and/or setting the stator vanes at a position for lower compression ratios, also results in inefficiencies. Thus, there is the dilemma that if the turboprop engine is designed to optimize operation at cruise, it suffers certain inefficiencies during the take-off and climb mode.
One prior art approach to alleviate this problem is to combine the turboprop with a turbofan engine so that there is both a propeller and a fan. This concept was discussed briefly in a paper entitled "Advanced Turboprop Transport Development - A Perspective", presented by Murray A Booth at the 13th Congress of the International Counsel of Aeronautical Sciences, August 22-27, 1982, Seattle, Washington.
A search of the patent literature disclosed a number of patents showing various devices to divert or bypass flow in jet engines. Although these in general do not relate directly to turboprop engines, these patents are cited herein to insure that the applicants are complying with their duty to make a full disclosure of all prior art of any possible relevance. These patents are as follows:
U.S. Pat. No. 2,978,865, Pierce, shows a device for mixing the exhaust of a turbofan engine.
U.S. Pat. No. 3,117,750, Snell, shows an aircraft engine having devices for selectively diverting the airflow.
U.S. Pat. No. 3,548,597, Etessam, discloses an aircraft gas turbine engine having a supplementary compressor driven by the exhaust gasses to precompress air that is fed to the main compressor. The embodiments shown are stated to be applicable to turboprop engines, hellicopter engines, jet engines, supersonic power plants incorporating the after-burners and power plants for vertical-take-off aircraft.
U.S. Pat. No. 3,635,029, Menloux, discloses a composite gas turbine ramjet engine where there is an outer through duct for ramjet operation.
U.S. Pat. No. 3,638,428, Shipley et al, discloses a fan jet engine having a low pressure and a high pressure compressor section. There is a device to bypass air to the fan duct to optimize flow to the high pressure compressor section.
U.S. Pat. No. 3,913,321, Snell, shows a gas turbine engine where there is a forward low pressure section and a year high pressure section. The low pressure section has forward and rear portions. Flow from the forward portion can, in one operating mode, be bypassed around the second portion, and inlet air can be caused to bypass the forward portion and pass directly into the second portion.
U.S. Pat. No. 4,038,818, Snell, shows another arrangement having generally the same operating principles as the above mentioned Snell patent.
U.S. Pat. No. 4,043,121, Thomas et al, discloses a two spool engine where there is an auxiliary fan positioned at the periphery of the fan of the conventional turbofan engine. Variable vanes and nozzles are utilized to vary the bypass ratio.
U.S. Pat. No. 4,052,845, Tumavicus, shows a variable ratio bypass turbojet engine having a particular flapper valve mechanism to vary the bypass ratio. During cruise, there is a low bypass ratio so that the front fan air is conducted to the core engine. For take-off condition, the front fan air is discharged to atmosphere through a thrust nozzle and the auxiliary inlet air is directed to the core engine fan.
U.S. Pat. No. 4,054,030, Pederson, discloses a variable cycle gas turbine engine having specific, mechanical devices to selectively bypass air from the forward compressor stage.
U.S. Pat. No. 4,060,981, Hampton, shows a jet engine having co-annular ducts. There is a valve device positioned between the forward and rear compressor sections which permits both the inner and outer upstream duct flows to be directed to either the inner or outer downstream duct, depending upon the orientation of the vane. Alternatively, either or both of the fluid streams can be blocked off entirely
In U.S. Pat. No. 4,292,802, Snow, there is a fan jet engine having blocker door vanes disclosed in the bypass duct to selectively close off the bypass flow and thus increase the airflow into the second compressor stage.
In British Patent Specification No. 713,783, there is shown a jet engine where the fan air can be selectively diverted into the second compressor stage.
In addition to the patents noted above, there are a number of U.S. patents issued to Garry W. Klees, one of the co-inventors in the present application, these patents being:
U.S. Pat. Nos. 3,779,282; 3,792,584; 3,854,286; 3,938,328; and 4,085,583. These patents show various arrangements for air breathing gas turbine engines having a particular configuration for a flow bypass mechanism which has inner and outer annular portions. In one position, airflow into the outer annular portion can be passed through to an outer annular duct, and air passing into the inner annular portion is passed into an inner annular duct. In another position, the flow can be reversed, so that the outer annular flow can be diverted into the inner annular duct, and the inner annular flow diverted outwardly into the outer annular duct. This particular valving arrangement has been found to be quite advantageous and is incorporated into the preferred embodiment of the present invention which is described later herein.
In view of the foregoing, it is an object of the present invention to provide a turboprop engine and a method of operating the same, with a view to alleviating some of the above indicated problems relating particularly to turboprop engines.